A gas turbine engine comprises a compressor section, a combustor section and a turbine section which are arranged in this order. The compressor section may share a common rotor with the turbine section or both sections may comprise individual separated rotors. Alternatively the turbine section may comprise a rotor in a high pressure section and another rotor in a low pressure section. The compressor section is adapted to pressurize air and to deliver the pressurized air to the combustor section further downstream. In the combustor section, or simply combustor, the pressurized air is mixed with fuel and ignited for generating hot combustion gas which flows downstream to and through the turbine section. Thereby, the hot combustion gas drives the turbine section such that its rotor rotates. Thereby, energy contained in the combustion gas in form of pressure and velocity may be converted to mechanical energy which may for example be used for driving a generator to generate electrical energy.
In single shaft arrangements (having a common rotor) high pressure and low pressure parts are mechanically connected such that the turbine section drives the compressor section.
In twin-shaft arrangements (having two separate rotors) the low pressure turbine section is mechanically independent, i.e. drives only the output power shaft, and the high pressure turbine section drives the compressor section.
The turbine section may comprise a high pressure turbine part and a low pressure turbine part which are arranged adjacent to each other so that the low pressure turbine part is arranged downstream of the high pressure turbine part. For conversion of energy of the hot combustion gas the turbine section, in particular the high pressure turbine section, comprises nozzle guide vanes arranged in one or more rows, wherein each row is arranged in a particular axial position with respect to the rotor whose rotation axis runs along the axial direction. A turbine section may comprise one or more rows of nozzle guide vanes which are spaced apart in an axial direction. The rows of nozzle guide vanes belong to a stator part of the gas turbine and do not move during operation of the gas turbine.
Downstream of a row of nozzle guide vanes a row of rotor blades is arranged which are connected to the rotor shaft and rotate upon impingement of the hot combustion gas to their blade surfaces. The row of nozzle guide vanes upstream of the row of rotor blades are adapted to appropriately guide the hot combustion gas towards the rotor blades for optimizing the conversion of energy. Thereby, the nozzle guide vanes are subjected to especially high temperatures due to the hot combustion gas contacting the nozzle guide vanes and transferring thermal energy to the nozzle guide vanes. In particular, the nozzle guide vanes are considered to be the most critical components of the turbine section regarding thermal stress.
Hot combustion gases and the combustor exhaust can reach very high temperatures (above 1500° C.), in particular under transient operation conditions. Thereby, transient operation conditions may be operation conditions of the gas turbine, in which a gas turbine load is changing with time, in which fuel supply is changing with time and/or in which air supply to the combustor is changing with time, in particular very rapidly. In particular, transient operation conditions are different from steady state conditions.
The hot combustion gases may heat up the outer surface portion of the nozzle guide vane, while the nozzle guide vane may internally be cooled by for example air delivered from the compressor or steam delivered from a heat recovery system. Thereby, a steep temperature gradient may result between the inner and the outer portions of the nozzle guide vane. Thereby, the guide vane is strained to a high degree and may be the most likely engine component to fail, wherein this failure is primarily due to low cycle fatigue.
Thus, the temperature of the combustion gas to which the guide vanes are subjected must be limited in order to avoid damage of the nozzle guide vanes.
On the other hand gas turbine engines are inherently designed to operate at high gas temperatures, thereby improving their cycle efficiency. Thus, it is desired to operate a gas turbine at the maximum allowable temperature tolerable for the components, such as the nozzle guide vanes. Subjecting the components to a temperature above these limits may lead to permanent damage of these components. For example, a small increase of a guide vane temperature may reduce the service life by a substantial amount. To prevent turbine damage induced by excessive, prolonged combustor outlet gas temperature, the engine may be operated at a turbine peak temperature that is several degrees below the vanes critical lifecycle fatigue temperature. In conventional systems the turbine component may be protected by controlling engine operations parameters. Thereby, controlling may be based on a gas temperature measured at a point of the turbine section downstream of the first row of nozzle guide vanes.
The temperature of the combustion gases at the inlet of the high pressure turbine section may be too high to be measured directly, which presents problems in appropriately controlling the gas turbine.
The combustor exit temperature may also be referred to as turbine entry temperature. It may conventionally be controlled by for example limiting fuel flow delivered to the combustor. It may be derived or estimated from a temperature of the hot gases downstream of one or more turbine sections after energy has been extracted from the hot gas and the gas temperature. This temperature further downstream may correspondingly be reduced to a suitable level which may be practically measured.
Thereby, the combustion gas temperature may be measured by a plurality of thermocouples disposed either at the outlet of the turbine section or between a high pressure and a low pressure turbine section. In any case it is currently not possible to accurately measure the temperature of the combustion gas exiting the combustor. Thereby, appropriately protecting components of the turbine sections may be difficult.
While in a steady state operation condition the gas temperature measured further downstream the first row of nozzle guide vanes may appropriately be used for estimating the actual combustor exited temperature, this may be difficult during transient operation. In particular, during engine transients each engine acceleration and/or deceleration may induce a cycle of thermal stress especially to the nozzle guide vanes. Further, during these transients the actual combustor exit temperature may not be appropriately estimated based on temperature measurements further downstream. Especially during these transients the nozzle guide vane may be subjected to temperatures exceeding their limiting temperatures.
In particular, the temperature of the gas measured downstream the first row of nozzle guide vanes may not reflect the true combustor exit temperature, because the thermocouple used for measurements probing at the turbine exit may be constructed for accuracy and durability but not for quick response. Thereby, the thermocouple probe construction may result in a lag with a relatively slow response as compared to that of the critical turbine hardware. Using a gas turbine engine capable of full load acceptance in just over a couple of seconds the transient combustion gas temperature may quickly increase. Although consideration of this temperature lag may not be critical for engine accelerations of long duration, the delay may become most significant when attempting to accurately compensate for thermocouple dynamics during rapid accelerations of short duration.
Document U.S. Pat. No. 6,167,690 discloses a control system, wherein a turbine inlet temperature is derived as a function of a turbine outlet temperature.
The document “Investigation of non-linear numerical mathematical model of a multiple shaft gas turbine unit”, by SooYong Kim and Valeri P. Kovalesky, KSME International Journal, Volume 17, No. 12, pages 2087-2098, 2003, discloses a mathematical model to calculate characteristics of a multi-shaft gas turbine, wherein energy balances are applied.
There may be a need for a method of determining an exit temperature of a combustor of a gas turbine. Further, there may be a need for a method of controlling a gas turbine, in order to operate the gas turbine in the highest possible temperature range without deteriorating or damaging components of the gas turbine. In particular, there may be a need for a method of controlling a gas turbine, when the gas turbine is operated under temporary changing, i.e. transient, operation conditions.